Purge system for rocket propellant system



Aug. 28, 1962 R. N. ABILD ETAL 3,050,936

PURGE SYSTEM FOR ROCKET PROPELLANT SYSTEM Filed June 30, 1959 5 Sheets-Sheet 1 www |||I|| l- 055A 1 1:4 1|: Rm Numa/ m Il 1MM. www Nb n mHm. JOU wwaw A Hm? mswco A fl WM5 W/ .w

N M NSM l QM, c WM@ W Aug. 28, 1962 R. N. ABILD ErAL PURGE SYSTEM FOR ROCKET PROPELLANT SYSTEM Filed June 50, 1959 I5 Sheets-Sheet 2 /N VEN T0195 ROBERT NAB/LD CLAUDE 0. BPODERS HUGH S. C19/M' JAMES O. GALLANT @W .77 @9V/- AT TORNEY United States Patent Office 3,050,936. Patented Aug'. 28, 1952 PURGE SYSTEM FR RUCKET PRPELLANT SYSEEM Robert N. Abild, New Britain, Claude 0. Bruders, Simsbury, and Hugh S. Crim, Glastonbury, Conn., and James O. Gallant, Rehoboth, Mass., assignors to United Aircraft Corporation, East Hartford, Conn., a corpo ration of Delaware Filed .lune 30, 1959, Ser. No. 824,1;353 S Claims. (Cl. dll- 356) This invention relates to liquid rocket engines, more particularly to the propellant llow and control system for one of the stages of a rocket vehicle,

An object of this invention is to provide an improved rocket engine propellant ilow and control system.

Another object of the invention is to provide an improved purge control system for a rocket engine propellant flow and control system.

Another object of the invention is to provide a purge control system for la rocket engine propellant flow and control system which requires no separate action other than moving a power lever, and in `which the quantity of purge is set without resort .to a timing device.

Another object of the invention is to provide a purge control system for a dual propellant rocket engine in which there is no path where particles of the two propellants could come into contact from back-flow to cause an explosion.

Another object of the invention is to provide a purge control system for a rocket engine propellant flow and control system which purges the propellant conduits downstream of the propellant valves, as well as the combustion chamber, as soon as a signal is given to Start the engine, and -which purges the conduits and the combustion chamber when the engine is shut down.

Still another object of the invention is to provide a purge control system for a dual propellant rocket engine which, when a power lever is moved to start Athe engine, empties a rst pair of accumulators to purge both propellant conduits downstream of the propellant valves, as well as the combustion chamber, and at the time charges a second pair of accumulators, and which, when the power lever is moved to stop the engine, empties the second pair of accumulators to purge the conduits and the combustion chamber `and at the same time charges the iirst pair of accumulators in preparation for a subsequent start.

Other objects and advantages will be apparent from the following specification and claims, and from the accompanying drawings which illustrate an embodiment of the invention.

ln the drawing:

FIG. l is a schematic diagram of the propellant flow and control system of our invention.

FIG. 2 is an enlarged section view of the propellant utilization valve.

FIG. 3 is an enlarged section view of the main propellant oXidizer valve.

FIG. 4 is an enlarged section View of the main thrust control.

FIG. 5 is an enlarged section view of the purge control and accumulators.

Referring to FIG. i of the drawing in detail, lll indicates a rocket thrust chamber comprising combustion chamber 12 and nozzle 14. Two propellants, one a fuel and the other `an oXidizer, are separately fed -to the chamber. The propellants preferably are hypergolic and will ignite spontaneously when mixed in the chamber.

The basic flow path for each propellant will first be traced and then the detailed elements connected with each path `will be discussed.

Fuel such as liquid hydrogen is contained in tank 16 and will ow from the tank through conduit lil to centrifugal pump Ztl. Fuel flows from the pump through conduit 22 and past check valve 24 to thrust chamber itl where itis admitted to jacket 26 surrounding the thrust chamber. The fuel flows through the jacket to collector 23 at the upstream end of the thrust chamber. From here the fuel liows through conduit 3@ to pump drive turbine 32, which has inlet guide vanes 34 and which is connected to pump 2d by shaft 36. The fuel then flows through conduit 38 and past main propellant fuel valve iti to manifold 42 from which it is injected through a plurality of openings into combustion chamber l2.

Oxidizer such as liquid iluorine is contained in tank dit which is connected by conduit 46 to centrifugal pump The pump is mounted on gear shaft 49 which carries gear 5ft which meshes with idler gear 52, which in turn meshes with gear 54E on shaft 3o. Oxidizer iiows from the pump `through conduit 56, past propellant utilization valve 53, and past main propellant oxidizer valve 6i) to manifold o?, from which it is injected through a plurality of openings into combustion chamber l2 for mixing with the fuel.

Between fuel tank lo and check valve 24%, conduit 18, the casing for centrifugal pump 2t) and the upstream end of conduit 22. are surrounded by cooling jacket 64. The jacket is connected to the top of tank lo and contains boil-off gases from the tank. The purpose of the jacket is to keep conduit ld `and pump 2d as cool as possible prior to a start and thus minimize the possibility of Ithe vaporizing of the liquid fuel therein. Vent valve on is provided on the casing of pump 2d and normally is spring loaded closed. The vent valve permits the escape of boil-off vapors from the jacket and thus assures flow through the jacket. Ilhis prevents the collection of bubbles along the inner wall of the jacket and otherwise aids the cooling function of the jacket. Boilolf vapors escaping through the vent valve pass through pipe 58 to pipe 7n through which they are dumped overboard.

Vapor vent valve .72 is mounted in conduit 22 between pump Ztl and check valve 24. This valve normally is spring loaded open to allow fuel vapor in the line upstream of the check valve to flow through pipe '74 to vent pipe 7d. The vapor vent valve will close upon a build-up of fuel pressure upstream of check valve 24.

Bypass conduit 76 is provided in the fuel flow path between conduits 3d and 38 and allows fuel to be bypassed around pump drive turbine 32. Butterfly valve 7S is located in conduit 76 and the position of the valve determines the relative quantity of fuel flowing through turbine 32. When the valve is closed all of the fuel passes through the turbine. As the valve is opened a continuously decreasing proportion of the fuel delivered by pump 2li ilows through the turbine, the remainder flowing through conduit 76. The position of butterily valve '7a is controlled by a main thrust control to be described below.

Pressure vent valve titl is located in fuel conduit 30 between bypass conduit 7o and turbine 32 and normally is spring loaded closed. The valve prevents a fuel pressure build-up beyond structural limits in conduit 30 by venting fuel vapor through pipe 82 to pipe 70.

Main propellant fuel valve lll is connected by links 84 to stem 85 of piston $3 in valve actuator 90. Piston 88 normally is loaded by spring 92 to the left in actuator housing 94 to close the `fuel valve. The actuator will be pressurized to open lthe valve when the rocket engine is started as will be described below.

Between oxidizer tank Lid and combustion chamber 12, conduit 4o, the casing for centrifugal pump 43 and conduit 56 are surrounded by jacket 96. The jacket is connected by pipe 98 to helium tank 188, the helium being admitted -to the jacket under high pressure to prevent any leakage of the liquid iluorine oxidizer out of the system. Pressure regulator 182 maintains a constant pressure in jacket 96 regardless of variations in tank pressure. ln 4the event that liquid oxygen is used as the oxidizer, the jacket would be connected to the top of tank 44 similarly to the connection lfor fuel cooling jacket 64 so that boil-oir" gases `from the tank could -be used to cool the oXidizer system.

Propellant utilization valve 58 is mounted on struts 184 within oxidizer conduit 56 and, `as shown in FIG. 2, includes stationary housing 106 having guide 188 projecting therefrom in an upstream direction. Bullet 116i is the movable portion of the valve and cooperative with restriction 112 to dene valve area. Stem 111i projects in a downstream `direction from .the bullet and terminates in piston 116. The stem is supported by the end of guide 188 and the piston is supported by the walls of chamber 118 within housing 186. Spring 128 surrounds stem 114 between the flanged end of guide 188 and piston 116. Bellows 122 surrounds guide 188 and connects housing 106 and bullet 118. Spring tends to move bullet 11@` to the right in an opening direction and `this loading is opposed by a variable pressure admitted to the interior of Ibellows 122 through passage 124 in lower mounting strut 184.

The pressure in passage 124 is regulated by frapper valve 126, FIG. l, which is connected at one end to a pair of opposed bellows, 128 and 13s?. The pressure loading of the bellows rotates the apper valve about fulcrum 132 to vary the area of nozzle 134. Bellows 128 is connected by line 136 to pressure transmitter 138 at the bottom of oxidizer tank 44 and bellows 130 is connected by line 148 to pressure transmitter 142 at the bottom of fuel tank 16.

Actuating fluid -for the system is ted through pipe 144 from helium tank 108 as will be explained below, restriction 146 being located in the pipe immediately upstream of passage 124. By virtue of the structure, changes in the pressure loading between bellows 128 and 136 rotate apper valve 126 to vary the area of nozzle 184 and control the pressure within bellows 122 to adjust the position of bullet 110 accordingly.

Main propellant oxidizer valve 61D is located in the oxdizer conduit `downstream of propellant utilization valve 58. The oXidizer valve is mounted on struts 148 `as shown in FIG. 3, and is comprised of streamlined shell 150 which is movable longitudinally in the conduit, the flange on nose section 152 of the shell `contacting seat 154 when the valve is closed. The shell has slots 156 through which struts 148 extend and which permit longitudinal sliding of the valve structure.v Bearing 158 is connected lto the inner ends of the mounting struts and serves as a `guide 'for shaft 160 connected to partition 162. Bellows 164 is connected to the bearing structure and the partition and denes chamber 166 to which an actuating pressure is admitted through passage 168 in lower mounting strut 148 from helium tank 11)@ when the rocket is started.

Spring 170 is mounted between plate 172 and end wall 174 and normally serves to locate the valve against seat 154. Nose section 152 is loosely connected to end wall 174 by hub 176 to permit self-alignment of the nose section with .the seat when the valve is closed. Spring 178 surrounding the hub holds the nose section in position against 4the end plate.

The thrust of the rocket engine is governed by main thrust control 188` through regulation of butterfly valve 78 in bypass conduit 76. The thrust control is shown in detail in FIG. 4, 4and includes power lever 182 which is mounted on and rotates shaft 184. A pair of cams, 186 'and 188, are also mounted on the shaft and are rotated when the power lever is rotated. The surface ot `cam 186 has a ydual cam track, one track being in contact with needle valve 19@ which controls the admission of helium yfrom pipe 192, connected to helium tank 100, to pipe 194. Helium flows through pipe 194 into pipe 196, FIG. l, `from which a signal pressure is introduced to main propellant fuel valve housing 94, to bellows 122 in propellant utilization valve 58, and to bellows 164 in main propellant oxidizer valve 68. Pipe 196 also is connected to purge control 198, the operation of which will be described below. Pressure regulator 280 in pipe 192 maintains a constant helium pressure in the pipes.

The other track on cam 186 is in contact with needle valve 262 which controls the admission of fuel from conduit 22 through pipe 284, restriction 206 and passage 288 to nozzle 218. The area of nozzle 210 is controlled by `ilapper valve 212 which pivots about fulcrum 214 on flexible partition 216 and the right end of which is connected to a pair of opposed bellows, 218 and 220. Bellows 218 is evacuated Iwhile the interior of bellows 220 is connected by passage 222 .and pipe 224 to combustion chamber 12. The combustion chamber pressure within bellows 228 tends to rotate the apper valve about fulcrum 214 in a clockwise direction.

The midportion of iiapper valve 212 is loaded by a power lever imparted loading which tends to rotate the tlapper valve in a counterclockwise direction about its fulcrum. The power lever loading is imparted by cam 188 which contacts follower 226 on guide piston 228 which is guided in bushing 288. Spring 232 is mounted between guide piston 22S and plate 234 connected to the ilapper valve. Any diiierence between the input loading to the ilapper valve from cam 188 and the absolute combustion chamber pressure input results in rotation of the apper valve and variation of the area of nozzle 210.

When the area of nozzle 218 is varied the pressure in chamber 236, connected to passage 208 by passage 238, also is varied. 4Power piston 240 is located in chamber 236, the piston being loaded in an upward direction by spring 242. Piston stern 244 is connected by link 246, FIG. l, to butterfly valve '78 in bypass conduit 76.

A follow-up for the servo system is provided by lever 248 which is connected :to stem 244 at pivot 258 and which is connected at its left end to xed pivot 252. Spring 254 is mounted between the right end of follow-up levers 248 and the left end of iiapper valve 212. By means of the lever and the spring, motion of power piston 240 results in the restoring of -the flapper valve to a null position.

Pump overspeed protection is provided by bellows 256 which is internally connected by pipe 258 to fuel conduit 22 near the. discharge fuel pump 20. Should fuel pump speed tend to exceed a predetermined value, the pressure in conduit 22 will increase and bellows 256 will expand. Expansion of the bellows will raise bell housing 260 against the force of spring 262. Abutment 264 on the bell housing will Contact ilapper valve 212 and the apper valve will be rotated about fulcrum 214 in a clockwise direction to close nozzle 218. The resulting pressure build-up in chamber 236 will lower piston 240 and open butterfly valve 78. This will bypass more fuel, reducing the quantity flowing through turbine 32, and thus reduce pump speed.

The purge system for the rocket engine includes a pair of accumulators 266 and 268. As shown in FIG. 5, accumulator 266 is divided by Wall 270i into chambers 272 and 274 and accumulator 268 is similarly divided by wall 276 into chambers 278 and 280. The chambers in each of the accumulators are of different size depending upon the volume requirements for the individual systems connected thereto.

Purge control 198 contains four chambers, 282, 284, 286 and 288. The large diameter lower portion of each chamber contains a bellows 298 and a bellows end plate 291 which are loaded in a downward direction by a spring 292. A rod 294 extends upward from each bellows end plate and has a double-faced valve 296 attached to its upper end. The upper face of each valve cooperates with an upper seat 298 and the lower face of each valve cooperates with a lower seat 308 in the control casing. The elements in only one of the chambers have been speci cally identified in order to avoid unnecessary multiplication of reference numerals on the drawing. All of the elements, however, operate in the same direction at the same time, with all of the valves being moved in a downward direction by the springs to position the valves on lower seats 388 or with the valves being pressure loaded in an upward direction to position the valves on upper seats 298 as shown. Chambers 282 and 284 and their bellows elements are associated with the operation of accumulator 266i, while chambers 286 and 288 and their bellows elements are associated with the operation of accumulator 268.

Pipe 382 connects helium tank 108 to purge control 198 and includes pressure regulator 384 FIG. l, for maintaining a constant pressure in the purge system. Inside the purge control the feed pipe is connected by branch pipe 386 to the top of chamber 288 at the valve upper seat, and by branch pipe 368 to the top of chamber 286 at the valve upper seat. Pipe 302 also is connected to the reduced section midportion of chambers 28dand 282 immediately below the valvel lower seat.

The upper portion of chamber 282 is connected by pipe 318 to accumulator chamber 272; the upper portion of chamber 284 is connected by pipe: 312 to accumulator chamber 27d; and the upper portion of chamber 286 is connected by pipe 314 to accumulator chamber 27 8 and the upper portion of chamber 288 is connected by pipe 316 to accumulator chamber 288. The lower portion of each of the chambers is connected by branch pipe 318 to pipe 196 connected by pipe 194, FIG. l, to main thrust control 180.

Pipe 328 connects the reduced section midportion of chamber 288 immediately below valve lower seat 380 to check valve 322, FIG. l, located in oxidizer conduit 56 immediately downstream of oxidizer valve 60. The check valve normally is loaded by spring 324- to close pipe 328. Branch pipe 326 connects the top of chamber 284 at upper seat 298 to pipe 32u and the check valve.

Pipe 328 connects the reduced section midportion of chamber 286 immediately below the valve lower seat 380 to check valve 338, HG. l, located in fuel conduit 38 immediately downstream of fuel valve all. The check valve normally is loaded by spring 332 to close pipe 328. Branch pipe 334 connects the top of chamber 282 at the upper seat 298 to pipe 328 and the check valve..

As shown in FIGS. l and 4, between purge control 198 and check valve 330, pipe 328 is connected by branch pipe 336 to main thrust control 180. Check valve 338 normally is loaded by spring 348 to close branch pipe 336, but when the pressure in the branch pipe is sufficiently high to open the check valve, connection is afforded by passage 222 between the branch pipe and combustion chamber pressure sensing bellows 228.

Operation In order to operate the rocket engine, both propellant tanks must be lled with cryogenic propellants in a liquid state. The liquid fuel in tank 16 will lill conduit 18, the casing for centrifugal pump 2d, and conduit 22 as far as check valve 2d. As the tank and the conduits are being filled fuel vapors can escape through vapor vent valve '72 to be vented overboard through pipe 74 and pipe 70, but as the pressure of the liquid fuel in conduit 22 increases the vapor vent valve will close. A certain amount of liquid fuel will flow through check valve 24 and is changed to a vapor state since there is no cooling jacket for the fuel conduit downstream of the check valve. This vaporized fuel will lill the remainder of conduit 22, thrust chamber jacket 26, collector 28, conduit 38, bypass conduit 76, and conduit 38 as far as main propellant fuel valve Alti which will be closed. Any tendency of the vaporized fuel to build up excessive pressure results in the opening of pressure vent valve 8u in conduit 38 and the venting overboard through pipes 82 and 78 of the excess vapor.

As oXidizer tank 44 is filled the. liquid oxidizer will ll conduit d6, the casing for centrifugal pump 48, and conduit 56 as far as main propellant oxidizer valve 6? which will be closed. Propellant utilization valve 58 upstream from oxidizer valve 68 is an an intermediate position prior to actual operation of the engine.

Helium from tank 168 will be admitted through pipe 192 to needle valve 198 in main thrust control 180. The needle valve is closed at this time by virtue of the position of power lever 182 in its extreme counterclockwise, or 0E, position. Helium also is admitted from the tank through pipe 382 to purge control 198. Prior to start of the engine the springs 292 in purge control charnbers 282, 284, 286 and 288 have moved the bellows end plates 2% and connected valves 296 in each chamber downward so that the lower face of each valve is in contact with lower seat 308 for each chamber. As a result helium from pipe 302 will flow through branch pipe 306, the upper portion of chamber 288 and pipe 316 to accumulator chamber 286, and also through branch pipe 308, the upper portion of chamber 286 and pipe 514 to accumulator chamber 278. The result is that both chambers in purge accumulator 268 are charged. Helium from line 302 also is admitted to the midportion of chambers 284 and 282, but the helium is dead-ended here because of the position of the valves 296 on the seats 300.

To start the engine, power lever 282 is advanced in a clockwise direction to its low power position. Rotation of the power lever rotates cam 186 to permit needle valve 19t? to be opened by the helium pressure in pipe 192. Helium then is admitted through pipe 194 to distribution pipe 196. The right-hand branch of the distribution pipe conducts helium to branch pipe 318 and the lower portion of each of purge control chambers 282, 284, 286 and 288. This pressure will tact on the outside of the bellows 290 and move the bellows end plates and attached valves upward until each valve is seated on upper seat 298. The raising of the valves in chamber 282 and 28dI permits helium from pipe 302 to flow through the upper portion of chamber 282 and pipe 318 to accumulator chamber 272, and through the upper portion of chamber 28d and pipe 312 to accumulator chamber 274. This charges purge accumulator 266.

The raising of the valve in chamber 286 permits helium from accumulator chamber 278 to ilow through pipe 314, the upper portion of chamber 286 and pipe "328 to fuel conduit check valve 330. The helium pressure will open the check valve and helium will be discharged through conduit 38 and manifold 42 into combustion chamber 12, thus purging the fuel conduit downstream of fuel valve 4l). Helium in pipe 328 also will enter branch pipe 336 and the pressure of the gas will open check valve 338 in main thrust control 180 to admit helium to pipe 22d, connected to combustion chamber 12. This will purge the combustion chamber pressure line and the combustion chamber. The raising of the valve in chamber 288 permits helium from accumulator chamber 286 to flow through pipes 316, the upper portion of chamber 288 and pipe 320 to oXidizer conduit check valve 322. The helium pressure will open the check valve and helium will be discharged through conduit 56 and manifold 62 into combustion chamber 12, thus purging the oxidizer conduit downstream of oxidizer valve 6l). Both check valves '-330 and 322 will close when the pressure of the helium from accumulator 268 drops below the spring loading on the check valves.

The left-hand branch of distribution pipe 196 admits helium pressure to fuel valve actuator housing 94 to move piston 88 to the right and open fuel valve 40, and admits helium pressure to oxidizer valve bellows 164 to move shell to the left and open oxidizer valve 60. Helium also is admitted from the distribution pipe to utilization valve bellows 122 and associated flapper valve nozzle 134.

Rotation of the power lever and cam 186 permits the fuel pressure in pipe 204 to open needle Valve 282. Fuel then Will be admitted through restriction `206 to chamber 236 and to flapper valve nozzle 218. The servo system for regulating the position of bypass butterfly valve 78 thus is activated and is ready for thrust controlling operation. The rotation of the power lever also rotates cam 188 to load iiapper valve 212 through follower 226, guide piston 228 and spring 232. A selected thrust input signal to the apper valve thus is supplied which tends to rotate the apper valve in a counterclockwise direction about fulcrurn 214 and increase the area of nozzle 212. This prevents a pressure build-up in chamber 236 and spring 20E-2 holds piston 24@ in its upper position to maintain butterfly valve 78 closed. In this position, thrust control valve 78 provides the maximum flow of fuel vapor through the driving turbine, with the result that the maximum available power is applied to the pumps for starting. The butterfly valve will stay closed until the combustion chamber pressure sensed by bellows 220 reaches a value sufficient to rotate the flapper valve in a clockwise direction, causing an increase of fuel pressure in chamber 236 to overcome the load of spring 242, and move piston 2li@ downward.

When fuel valve 40 land oxidizer valve 60 have been opened by helium pressure, tfuel vapor and liquid oxidizer will start to iiow at tank pressure to combustion chamber 12 where they will ignite spontaneously upon mixture with each other.

The heat of the combustion gases lin thrust chamber increases the temperature of jacket 26 surrounding the thrust chamber. This increases the temperature of the fuel vapor therein, increasing its energy level. This increase in energy permits turbine 32 to accelerate fuel pump and oxidizer pump 48, increasing the propellant pressures in conduits 22, 30, 38 and 56. As the speed of the propellant pumps increases, a continuously increasing quantity of fuel and oxidizer is delivered to the combustion chamber. This will increase the combustion level in the combustion chamber to generate more heat and to further heat the fuel vapor in jacket 26. The fuel vapors in turn will drive turbine 32 faster to further increase propellant pump speed and propellant flow.

With an increased flow of propellants to combustion chamber 12, the pressure sensed by bellows 220 in the main thrust control is increased yand flapper valve 212 is rotated in a clockwise direction. This will increase the pressure in chamber 2136 to lower piston 240 and open bypass valve 78. At the point where the forces acting upon dapper valve 212 are in equilibrium the quantity of vaporized fuel flowing to turbine 32 is just suicient to drive the propellant pumps at a delivery rate producing the selected thrust and the excess of the fuel is bypassed through conduit 76. A change in power lever position to increase or decrease the selected thrust results in an input signal to the flapper valve which produces an appropriate change in bypass valve opening so that the propellant pumps are driven at the speed producing the required thrust. If the speed of the pumps should tend to exceed a predetermined limit, the resulting pressure build-up bellows 256 will raise abutment 264 to contact flapper valve 212 and close nozzle 210. The pressure in chamber .236 will be increased to drive piston 240 downward and open bypass valve 78. This will reduce the quantity of fuel owing to the pump drive turbine, and hence reduce the speed of the propellant pumps.

The position of propellant utilization valve 58 is adjusted in response to a signal derived from each propellant tank and sensed by bellows 128 and 130. The purpose of the valve is to insure that the two propellants are used in their correct proportions, and both propellant tanks are emptied at the same time. A difference in the pressure forces of the bellows will rotate flapper valve 126 to vary the area of nozzle 134i and vary the pressure within valve bellows 122. The position of bullet thus will be changed to control the flow area of conduit 56 and either increase or decrease oxidizer flow rate in proportion .to the fuel flow rate.

To shut down the engine, power lever 182 is moved to its olf position and cam 186 will close needle valves 194D and 282. The closing of needle valve 190 cuts off the helium signal pressure to distribution pipe 196. The helium pressure in the distribution pipe is vented through drain line 342 in the main thrust control, the interior of the control being connected to vent pipe 78.

When the signal pressure to the distribution pipe is cut off, spring 92 will close fuel valve 4@ and spring 17) will close oxidizer valve 68. Closing of the valves terminates the ilow of propellants to combustion chamber 12.

The venting of the signal pressure in distribution pipe 196 also will relieve the pressure in the lower portions of purge control chambers 282, 284, 286 and 288. The spring 292 in each chamber will force each valve downward to sit on lower seat 308. The lowering of the valve in chamber 282 results in helium from accumulator charnber 272 owing through Ipipe 310, the upper portion of chamber 282, branch pipe 334.1, and pipe 328 to fuel conduit check valve 330. The helium pressure will open -the check valve and purge the fuel conduit downstream of fuel Valve 40. Helium also is admitted from pipe 330 through branch pipe 336 to open check valve 338 in the main thrust control and purge combustion pressure pipe 224. The lowering of the valve in chamber 284 permits helium from accumulator chamber 274 to iiow through pipe 312, lthe upper portion of chamber 284, branch pipe 326, and pipe 32@ to oXidizer conduit check valve 322. The helium pressure will open the check valve and purge the oxidizer conduit downstream of oXidizer valve 69.

The lowering of the valves in each of chambers 286 and 288 permits the recharging of accumulator chambers 278 and 281) in purge accumulator 268.

After the engine has been shut down, propellant fiow will have been stopped by `the closing of the fuel and oxidizer valves, the pump drive turbine will stop, both propellant lines will have been purged, and purge accumulator 268 will be charged in anticipation of operation of the engine again.

It is to be understood that the invention is not limited to the specific embodiment herein illustrated and described but may be used in other ways without departure from its spirit as defined by the following claims.

We claim:

1. In a dual propellant rocket engine system having a supply conduit for each propellant, a combustion chamber to which the propellants iiow and means for starting and stopping said engine, a shutoff valve in each conduit, means actuated by said starting `and stopping means for purging said conduits downstream of said shutoff valves and said combustion chamber when said engine is started and means actuated by said starting and stopping means for purging said conduits downstream of said shutoff valves and said combustion chamber when said engine is stopped.

2. In a dual propellant rocket engine system having a supply conduit for each propellant, a combustion chamber to which the propellants iiow and means for starting and stopping said engine, a shutoc valve in each conduit, a pair .of purge accumulator means, means for discharging one accumulator means to purge said conduits downstream of said shutoff valves when said engine is started, means for simultaneously charging said other accumulator means, means for discharging said other accumulator means to purge said conduits downstream of said shutoff valves when said engine is stopped, and means for simultaneously charging said one accumulator means.

3. In a dual propellant rocket engine system having a supply conduit for each propellant, a combustion chamber to which the propellants flow and means for starting and stopping said engine, a shutoff valve in each conduit, a pair of purge accumulator means, means for discharging one accumulator means to purge said conduits downstream of said shutoff Valves and said combustion chamber when -said engine is started, means for simultaneously charging said other accumulator means, means for discharging said other accumulator means to purge said conduits downstream of said shutoff valves and said combustion chamber when said engine is stopped, and means for -simultaneously charging said one accumulator means.

4. In a dual propellant rocket engine system having a supply conduit for each propellant, a combustion cham- Iber to which the propellants flow and means for starting and stopping said engine, a shutoff valve in each conduit, a first pair and a second pair of purge accumulators, rst valve means controlling said first pair of purge accumulators and `second valve means controlling said second pair of purge accumulators, means for shifting said first valve means to discharge said first pair of purge accumulators to purge said conduits downstream of said shutoff valves and said combustion chamber when said engine is started, means for simultaneously shifting said second Valve means to charge said second pair of purge accumulators, means for shifting said second valve means to discharge said second pair of purge accumulators to purge said conduits downstream of said shutoff lValves and said combustion chamber when said engine is stopped, and means for simultaneously shifting said first valve means to charge said first pair of purge accumulators.

5. In a propellant iiow and control system for a rocket engine including a thrust chamber having a combustion chamber therein, control means for starting and stopping the engine, a fuel supply conduit and an oxidizer supply conduit to said combustion chamber, means for pumping said propellants through said conduits, means for heating the fuel in a portion of said fuel conduit, means for driving said pumping means by expansion of said heated fuel, means responsive yto combustion chamber pressure for variably controlling the Jfiow of heated fuel to said driving means, means actuated by said control means for purging a portion of said conduits and said pressure sensing means when said engine is started, and means actuated -rby said control means for purging a portion of said conduits and said pressure sensing means vwhen said engine is stopped.

6. In a propellant flow and control system for a rocket engine including a thrust chamber having a combustion chamber therein, a fuel supply conduit and an oxidizer supply conduit to said combustion chamber, means for pumping said propellants through said conduits, means for heating the fuel in a portion of said fuel conduit, means for driving said pumping means by expansion of said heated fuel, means responsive to combustion chamber pressure for variably controlling the flow of heated fuel to said driving means, two pairs of purge accumulators, means for discharging the first pair of accumulators to purge a portion of said conduits when said engine is started, means for simultaneously charging said second pair of accumulators, means for `discharging said second pair of accumulators to purge a portion of said conduits when said engine is stopped, and means for simultaneously charging said first pair of accumulators.

7. In a propellant liow and control system for a rocket engine including a thrust chamber having a combustion chamber therein, a fuel tank, conduit means connecting said tank and said combustion chamber, fuel pumping means in said conduit means, means for heating -fuel in a portion of said conduit means by heat generated in said combustion chamber, driving means operatively connected with said pumping means, means for driving said driving means by expansion of said heated fuel, means for regulating the quantity of heated fuel driving said driving means, valve means controlling the admission of fuel to said combustion chamber, an oXidiZer tank, conduit means connecting said tank and said combustion chamber, oXidizer pumping means in said conduit means, valve means controlling the admission of oXidiZer to said combustion chamber, means for automatically controlling said fuel regulating means in accordance with combustion chamber pressure, two pairs of purge accumulators, means for discharging the first pair o-f accumulators to purge said conduit means downstream of said admission controlling valve means when said engine is started, means for simultaneously charging said second pair `of accumulators, means for discharging said second pair of accumulators to purge said conduits downstream of said admission controlling valve means when said engine is stopped, and means for simultaneously charging said :first pair of accumulators.

8l A propellant ow and control system for a rocket engine, said system including a fuel tank and an oxidizer Itank each connected by conduit means to the combustion chamber in a thrust chamber, said fuel conduit means including a pump therein, a check valve downstream of said pump, a vapor vent valve between said pump and said check valve, a jacket downstream ,of said check valve and surrounding said thrust chamber, a drive turbine downstream of said jacket adapted to be driven by the expansion of lfuel heated in said jacket and operatively connected with said fuel pump, a `bypass conduit around said turbine, a pressure vent valve between said turbine and the inlet for said bypass conduit, a normally closed lfuel valve downstream of said turbine and the outlet for said bypass conduit, and 4a mainfold through which fuel is injected into said combustion chamber, said oXidiZer conduit means including a pump .therein operatively connected with said drive turbine, a propellant utilization valve for trimming oxidizer `fiow located downstream of said oxidizer pump and responsive to the liquid level in said fuel and oXidiZer tanks, a normally closed oxidizer valve downstream of said utilization valve, and a manifold through which oxidizer is injected into said combustion chamber, valve means in said lbypass conduit, means for opening said fuel and oXidizer Valves Ito permit propel lant ow to said combustion chamber, a main thrust control for automatically regulating the position of said bypass valve means to determine the flows of heated fuel through said turbine and said bypass conduit, a first pair and a second pair of purge accumulators, first valve means controlling said first pair of purge accumulators and second valve means controlling said second pair of purge accumulators, means for shifting said rst valve means to discharge said first pair of purge accumulators to purge said fuel and oXidizer conduits downstream of said normally closed valves when said engine is s-tarted, means for simultaneously shifting said second valve means to charge said second pair of purge accumulators, means for shifting said second valve means to discharge said second pair of purge accumulators to purge said fuel and oxidizer conduits downstream of said normally closed valves when said engine is stopped, and means for simultaneously shifting said rst valve means to charge said first pair of purge accumulators.

References {lited in the file of this patent UNITED STATES PATENTS Harby Sept. 27, 1949 Goddard Dec. 5, 1950 OTHER REFERENCES 

